1. Field of the Art
The present invention relates generally to a vortex flow field and the apparatus and method to produce and sustain it and more particularly to a hybrid rocket engine and a method of propelling a rocket utilizing such vortex flow field. The flow field in accordance with the present invention is capable of providing separate regions or zones within and among one or more flowing fluids contained within a common chamber, without the need for diaphragms or other physical separators or barriers. It is evident and believed that the flow field of the present invention has utility to a wide range of applications. One general field of application is that of combustion chambers, and more particularly, that of combustion chambers and methods for rocket engines or the like and hybrid rocket propulsion systems. A combustion chamber and method in accordance with one embodiment of the present invention utilizes the unique flow field of the present invention to improve hybrid rocket fuel regression and increase mixing length in a rocket or other engine. Another embodiment is in the form of liquid rocket engine to prevent hot combustion products from contacting the chamber wall.
2. Description of the Prior Art
Virtually countless applications exist for a flow field which is compact and is capable of providing one or more separate regions or zones of flowing fluids within a container, without substantial mixing and without the need for any physical barrier or other separators between such regions or zones. With such a flow field, a chemical reaction, such as combustion, can be induced to incur in one region or zone while a separate fluid or process occupies another region or zone.
Many devices depend upon vortex flows for their successful operation, such as combustion chambers, cyclone separators, classifiers and the like that are in common use. All of these devices introduce swirling flow at one end of a passageway in which the flow follows a generally helical path to exit at the opposite end. Such conventional vortex flows do not achieve zonal separation as does the unique flow field that is the subject of the present invention.
Although the flow field in accordance with the present invention has significant applications in a variety of fields, it has particular application to the field of rocket engines and in one embodiment, specifically to hybrid rocket engines. Hybrid rocket engines denote a class of rocket propulsion systems in which one propellant is in fluid form and the other propellant is in the form of a solid grain. Typically, the fluid propellant is the oxidizer and the solid grain is the fuel. The oxidizer such as liquid oxygen is sprayed into the combustion ports in the solid fuel grain and caused to ignite. The hot combustion products sustain the combustion process until either the oxidizer flow is shut off or the fuel grain is depleted. In virtually all contemporary hybrids of today the limiting design factor is the rate at which the solid fuel grain can be caused to burn. The burn rate, often expressed as regression rate, is the rate at which fuel can be induced to vaporize or ablate off the grain surface so it can participate in the combustion process and contribute to rocket thrust. Because the rate is typically slow, conventional hybrid fuel grains must be made with large exposed surface areas. This is accomplished by casting large open combustion ports in the grain. The large ports waste volume in the high pressure casing, so that a larger, heavier, and more expensive case is needed than would be required if the fuel grain combustion ports could be much smaller by means of a flow field which improves the regression rate.
In recent years, hybrid rockets have received increasing attention from the National Aeronautics and Space Administration (NASA) sectors, Department of Defense, industrial aerospace participants and research institutions because their unique operational characteristics are capable of providing safer, lower-cost avenues to space than conventional solid propellant and liquid bi-propellant rocket propulsion systems. For example, hybrid rocket engines can be easily throttled for thrust tailoring, to perform in-flight motor shutdown and restart and to incorporate non-destructive mission abort modes. Also, since fuel in a hybrid rocket engine is stored in the form of a solid grain, such engines require only half the feed system hardware of liquid bi-propellant engines. Still further, the commonly used butadiene-based solid grain fuels are benign and neither toxic nor hazardous for storage and transportation. The hybrid solid fuel grain is also not susceptible to cracks and imperfections that can lead to catastrophic failure in solid rocket motor propellant grains.
However, despite these benefits, classical hybrid rocket engines, in which the oxidizer gas is injected into the combustion chamber at the end opposite the exit nozzle and in a direction parallel to the solid fuel grain, have not yet found widespread use for either commercial or military applications. Reasons for this include the fact that they suffer from relatively slow solid fuel regression rates, low volumetric loading and relatively poor combustion efficiency. For example, polymeric hybrid fuels such as hydroxyl-terminated polybutadiene (HTPD) regress generally about an order of magnitude slower than solid rocket motor propellants. In an effort to overcome these lower regression rates, complex cross-sectional geometries of the hybrid solid grain fuel with large wetted surface areas are often employed to achieve a large mass of flow rate of pyrolyzed vapor from the fuel grain. It has been shown that a three to fourfold increase in fuel regression rate can result in significant cost reductions, simplified grain manufacturing and large reductions in rocket inert weight.
In addition to problems associated with the low regression (fuel burning) rates of hybrid engines, the short straight line travel of the pyrolyzed fuel grain vapor and oxidizer as they traverse the combustion region results in incomplete mixing. This often necessitates the use of secondary combustion chambers at the end of the fuel grain to complete the combustion process. These secondary chambers add length and weight to the overall design and have the additional disadvantage of serving as a potential source and location of combustion instability.
Furthermore, both conventional hybrids and solid rocket motors must provide insulation layers between the solid propellant grain and the high pressure casing wall. This is necessary to prevent the exposure of the casing to the high temperature combustion gases when the grain material has been burned away out to the casing and no longer provides protection. The insulation adds weight and cost to the motor.
Accordingly, there is a need in the art for a flow field, and a structure and method for producing and sustaining it, which provides separate regions or zones of flowing fluids within a chamber. There is also a need in the art for a combustion chamber and method utilizing such a flow field, and particularly a combustion chamber and method for a hybrid rocket engine, which significantly increases the regression rate of the solid fuel grain and the effective chamber length and mixing within the combustion chamber. There is also a need for a combustion chamber and method utilizing such a flow field that prevents the hot combustion products from reaching the chamber wall.
In accordance with the present invention, a fluid flow field, and a structure and method for producing and sustaining the field, has been designed. This flow field provides for separate regions or zones of flowing fluids within a chamber without the need for physical barriers or other separators and without substantial mixing between the regions or zones.
In a revolutionary departure from prior art the present invention introduces the incoming swirling flow concentric to the outlet passage and by this means establishes a new and unique flow field not here-to-fore known or described in literature, nor does it have any previously known existing physical embodiments beyond those defined and described herein. The flow field inherently divides into an outer upwardly flowing vortical helix along a chamber wall, an inner downward flowing vortical helix along the center region of the chamber, a converging flow field at the head end where the outer vortex transforms into the inner vortex, a converging flow field as the flow approaches the exit nozzle, and less well defined regions of velocities and pressure gradients elsewhere throughout the chamber.
The distinct regions can be controlled by chamber geometry, fluid injection parameters, external heat addition, and combustion or other chemical reactions to produce certain desired and specific results. These reactions include, but are not limited to, enhanced combustion of materials forming the chamber walls, limitation of combustion to occur in the center vortex only, combination of reactions at the wall, with subsequent separate and different reactions in the central vortex, and fluid distillation and liquid-vapor separation.
The flow field is produced by injecting flow tangentially into a cylindrical chamber which is substantially closed at one end and which has a converging outlet at the other end. In the preferred embodiment, the flow is introduced into the interior of the chamber near the outlet end of the chamber and in a direction which is substantially tangent to, or which creates a flow which is substantially tangent to, the inner wall of the chamber. This tangential injection causes the flow in the chamber to swirl and follow a spiral path up the inner wall of the chamber thereby establishing an annular vortex flow bounded by the inner wall of the chamber. When the spiral flow reaches the closed end of the chamber, the flow inherently translates inwardly to the center of the chamber to form the second vortex where the flow moves spirally away from the closed end, down the core of the chamber and out the chamber nozzle. This inner vortex or spiral flow through the center of the chamber rotates in the same direction as the outer vortex, but at a smaller radius and thus a greater angular velocity in accordance with the principle of angular momentum conservation. The result of the above is the establishment of a radial pressure gradient field throughout the chamber. One key feature of this pressure gradient field occurs at the exit nozzle. Specifically, pressure at the nozzle converging wall increases and pressure at the swirl axis decreases. Accordingly, injection flow at the periphery of the vortex near the outlet end and tangential to the outer vortex streamline cannot penetrate the pressure gradient that has formed by the inner vortex at the nozzle converging region. Thus, this incoming flow cannot flow toward the exit. Instead, it must take an alternate flow path to enter the lower pressure region in the center of the vortex flow approaching the converging nozzle section. This alternate path is up along the wall and then inward to the center vortex before flowing down and out the nozzle. Accordingly, as the inner vortex flow approaches the nozzle, it enters the converging section of the nozzle, thereby increasing the swirl or angular velocity and thus producing an enhanced radial pressure gradient that blocks the outflow of the fresh incoming stream.
The above-described flow field has several unique characteristics. First, the flow path of the injected fluid before reaching the outlet is quite long and highly convoluted. Thus, it provides an opportunity for intense and extensive mixing along the flow path, particularly in the core or inner vortex where the angular velocity of the swirl is greater. Secondly, the outer and inner vortexes are individually discrete. Thus, the fluid flow in the inner vortex does not mix significantly with the fluid flow in the outer vortex. This enables the inner vortex to support burning or other chemical reactions to some significant degree independent of the outer vortex. Because of this, materials such as propellant or other chemicals, can be added to the inner vortex by injection at the conjunction of the two vortices at the closed end of the chamber and cause combustion or other chemical reaction to occur and be sustained wholly in the inner vortex if so desired.
The ability to produce and sustain the above-described double vortex field flow has countless potential applications and several immediate practical applications. By way of example only, one immediate practical application of the flow field of the present invention is in the field of rocket propulsion.
In such an application, utilization of the flow field of the present invention facilitates a combustion chamber and method which provides dramatically increased regression rates of the fuel grain and increased mixing length and improved mixing within the combustion chamber. In a preferred embodiment and application, the present invention provides for a combustion chamber and method for use in a hybrid rocket engine.
In applying the double vortex flow field of the present invention to the preferred embodiment of a hybrid rocket engine, the flow is created by injecting one component of the combustion mixture (such as the oxidizer) into a generally cylindrical combustion chamber which is closed at one end and is provided with a converging outlet nozzle at the other end. By injecting the flow of oxidizer fluid in a direction circumferentially tangent to the inner wall, the fluid is caused to swirl and advance up the cylinder wall in a vortex pattern toward the closed end. When this outer vortex flow reaches the closed end, it moves radially toward the center of the chamber and continues to move in a swirling vortex along the middle or core of the chamber and out through the exit nozzle. If the inner walls of the chamber are hybrid fuel grain and the fuel grain/oxidizer combination is ignited, several unique and advantageous characteristics result. First, the high velocity outer vortex scrubs the burning fuel grain surface, causing enhanced heat transfer to the surface. Combustion near and on the surface is also able proceed because fresh oxidizer is carried directly to the surface by turbulent transport mechanisms in addition to the usual molecular diffusion process. Second, the vortex also sustains radial pressure and density gradients that cause hot, low density combustion products to be buoyed out of the combustion zone so their presence does not hinder the combustion process.
Third, because the flow path of the injected fluid (the oxidizer) to reach the outlet is very long and highly convoluted, it provides an opportunity for intense and extensive mixing and combustion with the fuel grain vapor, particularly in the core or inner vortex. Accordingly, in the above application, the outer vortex flow causes rapid burning of the fuel grain on the wall of the cylinder, and the inner vortex causes combustion to proceed rapidly, by providing intense mixing and combustion travel distance to allow combustion to reach completion, thereby achieving high combustion efficiency.
In general, the structure to produce the flow field of the present invention as well as the structure of the combustion chamber in accordance with the present invention includes a containment chamber with first and second ends which are sometimes referred to as head and tail ends or closed and nozzle ends. In the preferred embodiment, the container inner wall is covered with a solid fuel grain or a fuel source. The chamber is closed at one end and provided with an exit nozzle at its opposite end. One or more fluid (oxidizer) delivery ports are provided near the end of the container adjacent to the nozzle for the purpose of delivering an oxidizer (or other fluid) into the chamber tangentially to the inner surface of the fuel grain coating the inner wall of the chamber. After injection, the oxidizer swirls along the surface of the fuel grain toward the closed end, at which location it moves radially toward the center and then swirls in the form of the inner vortex toward the nozzle end of the cylinder.
The method aspect of producing the flow field of the present invention includes providing a cylindrical chamber with a closed head end and an opposing nozzle end and injecting a fluid tangentially to the inner wall. In the preferred application of the present invention the cylindrical chamber is a combustion cylinder with a closed end and a nozzle end and the inner surface of the chamber is provided with a fuel source. The injected fluid is an oxidizer. Upon injection of the oxidizer, the oxidizer/fuel mixture is ignited.
Accordingly, it is an object of the present invention to provide a flow field, and a structure and method for producing and sustaining such flow field which provides distinct and separate regions of flowing fluid within a chamber, without the use of physical barriers.
Another object of the present invention is to provide an improved combustion chamber and method utilizing the above-described double vortex flow field.
Another object of the present invention is to provide an improved combustion chamber and method utilizing the above flow field and to provide for increased fuel regression rates and increased travel distance and mixing to achieve complete combustion.
A further object of the present invention is to provide a hybrid rocket engine utilizing the above-described double vortex flow field.
A further object of the present invention is to provide a liquid rocket engine utilizing the above-described vortex flow field.
A still further object of the present invention is to provide an improved hybrid rocket propulsion system that facilitates and promotes high and uniform fuel grain regression rates so that small combustion ports can be used in the propellant solid grain.
Another object of the present invention is to provide a hybrid propulsion system that inherently cools the case walls whenever fuel is not present to insulate the wall from hot combustion products.
A more specific object of the present invention is to provide a hybrid rocket propulsion system that creates and uses a unique internal combustion vortex flow field to enhance grain regression rate and to increase the efficiency of the combustion process.
Another object of the present invention is to provide a combusting flow field that allows the use of a single grain port for the combustion process.
A further object of the present invention is to provide an injection means for the fluid propellant that induces the double vortex flow field in the grain combustion port.
A further object of the present invention is to provide a combustion process that inhibits combustion instability.
Another object of the present invention is to provide a double helix flow field in which an outer helix flows upwards along the grain surface inducing combustion, and an inner combustion helix flows down the port centerline and out the nozzle to produce thrust.
To achieve the foregoing and other objects and in accordance with the purpose of the present invention, a self-contained propulsion system is provided with a motor casing that houses a solid propellant grain. A first fluid propellant that will combust when in the presence of the solid propellant in the presence of an ignition source, is stored separately from the solid propellant in a fluid tank. A delivery means supplies at least a portion of the said fluid propellant in either liquid or gaseous state to the combustion port of the solid grain. An ignition means initiates combustion with the combustion port of the solid propellant grain. A fluid injection means that will cause the fluid propellant to enter the solid propellant grain case in such a manner as to form an upflowing helix along the surface of the combustion port in the solid propellant grain and then a downflowing helix along the centerline of the combustion port, said downflowing helix to eventually exit the chamber via the discharge nozzle.
The fluid propellant can be provided to the entrance to the fuel grain case by any of various common means, including delivery from pressurized tanks, or by pumps of suitable designs. The fluid can be either the liquid or gaseous state. Commonly the fluid propellant is the oxidant. In one embodiment, the oxidant is burned in a highly oxidizer-rich combustor (termed a xe2x80x9cpreburnerxe2x80x9d) and the resulting oxidizer-rich combustion products are used to drive a turbopump that pressurizes the liquid oxidizer for delivery to the preburner. After driving the turbine, the oxidizer-rich combustion products leave the turbine and flow to the injection ports of the fuel grain high pressure casing. The oxidant enters the fuel grain ports in a fluid phase that may be at high enough pressure to be supercritical. The injector elements are positioned and designed such that the injected flow develops the co-axial vortex flow field within the chamber in the manner that is the subject of this invention.
In another embodiment, the oxidizer in liquid state is carried in a high pressure tank. Pressurant is supplied by a conventional tank pressurization system well known to those acquainted with the profession. The liquid oxidant is expelled from the tank and delivered at high pressure to the injection ports of the fuel grain high pressure casing. The oxidant enters the casing in the liquid state and is quickly heated and vaporized as it enters the combustion port of the fuel grain. The injector elements, in concert with the cylindrical casing and cylindrical combustion port in the fuel grain, are designed to impart a strong swirl component to the injected oxidant flow. The swirl acts in the chamber to develop the co-axial vortex flow field that is a key aspect of this invention.
Typical oxidants are oxygen in liquid or gaseous form, inhibited red fuming nitric acid (IRFNA), hydrogen peroxide, nitrogen tetroxide and nitrous oxide.
The solid fuel grain may be of any suitable material. A preferred material used for hybrid fuel is hydroxyl-terminated polybutadine, a complex, rubber-like hydrocarbon formulation that is readily cast and cured at modest cost. Other fuels include paraffins and PMMA. Fillers such as aluminum powder and boron may be added to customize performance. The propellants chosen are not critical to the technology of this invention.
The above and other objects of the present invention will become apparent with reference to the drawings, the description of the preferred embodiment and the appended claims.